Segment for a turbine rotor stage

Abstract

A rotor stage ( 10 ) of a turbine engine includes a circumferential row of rotor segments ( 12 ), each including: first and second endwalls ( 14, 16 ) spaced apart radially, and a first and second sidewalls ( 18, 20 ) extending radially between the first and second endwalls ( 14, 16 ) and spaced apart circumferentially. The first and second endwalls ( 14, 16 ) and the first and second sidewalls ( 18, 20 ) define therewithin a flow passage ( 22 ) for hot gas. Circumferentially adjacent segments ( 12 a, 12 b ) mate along a respective split-line ( 24 ) extending along an interface between the first sidewall ( 18 ) of a first segment ( 12 a ) and the second sidewall ( 20 ) of a second circumferentially adjacent segment ( 12 b ). A composite airfoil structure ( 26 ) is thereby defined having a pressure sidewall ( 18 ) formed by the first sidewall ( 18 ) of the segment ( 12 a ) and a suction sidewall ( 20 ) formed by the second sidewall ( 20 ) of the second segment ( 12 b ). The first and second endwalls ( 14, 16 ) are respectively configured as a platform ( 14 ) and a tip shroud ( 16 ) of the segment ( 12 ).

1 . A rotor stage of a turbine engine, comprising:
 a circumferential row of rotor segments, each segment comprising:
 first and second endwalls extending in a circumferential direction and spaced apart in a radial direction in relation to an axis of the turbine engine, the first endwall configured as a platform and the second endwall configured as a tip shroud of the segment, and 
 first and second sidewalls spaced apart in the circumferential direction and extending radially between the first and second endwalls, 
 wherein the first and second endwalls and the first and second sidewalls define therewithin a flow passage for a hot gas, 
   wherein circumferentially adjacent segments mate along a respective split-line which extends along an interface between the first sidewall of a first segment and the second sidewall of a second circumferentially adjacent segment, to form composite airfoil structure which comprises:
 a pressure sidewall formed by the first sidewall of the first segment and a suction sidewall formed by the second sidewall of the second segment the pressure and suction sidewalls of the airfoil structure extending between a leading edge and a trailing edge of the airfoil structure. 
   
     
     
         2 . The rotor stage according to  claim 1 , wherein at least one of the segments is formed at least in part from a ceramic matrix composite material. 
     
     
         3 . The rotor stage according to  claim 2 , wherein the ceramic matrix composite material forms respective hot gas exposed surfaces of the first and second endwalls and the first and second sidewalls that define the flow passage of the segment. 
     
     
         4 . The rotor stage according to  claim 3 , wherein at least a portion of the segment comprises metallic substructure over which a skin made up of the ceramic matrix composite material is assembled to form the hot gas exposed surfaces of the segment. 
     
     
         5 . The rotor stage according to  claim 4 , wherein at least the first and second sidewalls are entirely formed of the ceramic matrix composite material. 
     
     
         6 . The rotor stage according to  claim 3 , wherein said hot gas exposed surfaces of the segment are formed by a continuous lay-up of the ceramic matrix composite material along an inner periphery of the segment which defines a boundary of a gas path volume of the flow path. 
     
     
         7 . The rotor stage according to  claim 1 , further comprising one or more stiffening beams extending radially outwardly from the tip shroud and running circumferentially along the tip shroud. 
     
     
         8 . The rotor stage according to  claim 1 , wherein the segment is attachable to a rotor disc via a root,
 wherein the split-line extends through the root,   wherein the root comprises a first root portion formed on the first endwall of the first segment and a second root portion formed on the first endwall of the second segment.   
     
     
         9 . The rotor stage according to  claim 1 , wherein either one the pressure sidewall or the suction sidewall of the airfoil structure is cutback from the trailing edge. 
     
     
         10 . The rotor stage according to  claim 1 , wherein the split-line extends along a mean camber line of the airfoil structure ( 26 ). 
     
     
         11 . The rotor stage according to  claim 1 , wherein the airfoil structure comprises an internal cavity defined between the pressure sidewall and the suction sidewall. 
     
     
         12 . The rotor stage according to  claim 11 , wherein the airfoil structure comprises a first gap at the leading edge and a second gap at the trailing edge, the first and second gaps being formed along a split-line interface of the pressure sidewall and the suction sidewall. 
     
     
         13 . The rotor stage according to  claim 12 , wherein the split-line is offset from a mean camber line of the airfoil structure toward the pressure sidewall or the suction sidewall of the airfoil structure, such that the first gap-and the second gap are correspondingly offset toward the pressure sidewall or the suction sidewalk of the airfoil structure. 
     
     
         14 . The rotor stage according to  claim 12 , wherein the split-line interface at the leading edge and/or at the trailing edge includes a ship-lapped interface. 
     
     
         15 . The rotor stage according to  claim 12 , wherein the internal cavity of the airfoil structure is pressurized by a fluid to maintain a positive outflow margin at the first and second gaps in relation to a hot gas flow external to the airfoil structure. 
     
     
         16 . The rotor stage according to  12 , further comprising radially extending coolant passages through the pressure sidewall and/or the suction sidewall of the airfoil structure, for conducting coolant between the first and second endwalls. 
     
     
         17 . The rotor stage according to  claim 16 , wherein the coolant passages are formed through the ceramic matrix composite material. 
     
     
         18 . The rotor stage according to  claim 16 , wherein the coolant passages are formed through a metallic substructure of the pressure sidewall and/or the suction sidewall. 
     
     
         19 . The rotor stage according to  claim 12 , wherein the first and second gaps are configured to allow hot gas ingestion into the internal cavity of the airfoil structure. 
     
     
         20 . A segment for a turbine rotor stage, comprising:
 first and second endwalls extending in a circumferential direction and spaced apart in a radial direction in relation to an axis of the turbine engine, the first endwall configured as a platform and the second endwall configured as a tip shroud of the segment, and   first and second sidewalls spaced apart in the circumferential direction and extending radially between the first and second endwalls,   wherein the first and second endwalls and the first and second sidewalls define therewithin a flow passage for a hot gas, and   wherein the respective segment is configured to mate with circumferentially adjacent segment on either side along a respective split line, such that each split-line-extends along an interface between one of the first or second sidewalls it of the respective segment and a corresponding other of the first or second sidewalls of the circumferentially adjacent segment on either side.